The invention relates to a measuring arrangement for determining the sun and earth orientation of a three-axis stabilized satellite and a corresponding process for determining the rotational velocity of the satellite and the deviation of the satellite as well as to a system for implementing attitude control maneuvers by using such a measuring arrangement and corresponding regulating processes.
A measuring arrangement of the initially mentioned type is known from H. Bittner, et al., "The Attitude Determination and Control Subsystem of the Intelsat V Spacecraft", Proceedings of AOCS Conference, Nordwijk, Oct. 3 to 6, 1977, ESA SP-128, November 1977. The satellites of this known type have a redundantly designed speed gyroscope package which measures in all three axes X, Y, Z of the satellite-fixed system of coordinates, as well as several sun sensors which are also designed in a redundant manner and have two fields of view. One field of view is centered about the negative Z-axis, which comprises one half of the XZ-plane, and has a specific width perpendicularly with respect to it (in the direction of the Y-axis). The other filled of view is centered about the positive X-axis, comprises one third of the XY-plane and also has a specific width perpendicularly to it. This sensor arrangement is used for tracking, in a measuring manner, rotations of the satellite body, that is, the components of the rotational speed vector of the satellite and, irrespectively of it, detecting the orientation of the satellite with respect to the sun by the determination of the sun vector pointing from the satellite to the center of the sun as soon as the sun is in the field of view of the sun sensors. The measuring results supplied by the sensors are used by the attitude control system of the satellite for commanding attitude changes or rotations in a targeted manner in order to be able to carry out the maneuvers required in the transfer orbit and in the final satellite orbit and to be able to correct occurring attitude deviations.
The speed gyroscopes used for the measuring, control and damping of rotational movements of the satellite are very complicated electromechanical precision devices, and are therefore correspondingly expensive and susceptible to disturbances. For meeting the reliability requirements under the extreme environmental conditions in space, particularly in the case of long-time missions, these devices must also be present in a redundant manner, which also increases the expenditures with respect to cost and weight. When one measuring axis is absent in a gyroscope package containing several speed gyroscopes, the remaining gyroscope axes can generally also no longer be used. Since the possibility of failure of a system rises with the number of functional elements, the use of a measuring system with a large number of electromechanical components which are susceptible to disturbances also has a higher risk of failure.
The invention is therefore based on the object of providing a measuring arrangement of the above-mentioned type which is less expensive and more reliable. By means of this measuring arrangement, it must be possible to carry out largely automatically on board of the satellite all maneuvers required in the transfer orbit as well as in the final satellite orbit, particularly in the geosynchronous orbit. Furthermore, if possible, the number of components which are susceptible to disturbances is to be reduced to a minimum.
The required maneuvers include: sun acquisition from any starting position; the earth acquisition from the sun orientation; alignment for the apogee maneuver, attitude stabilization in the desired orientation and the damping of movement during the maneuver period in the transfer orbit; as well as the earth and sun reacquisition in the geosynchronous orbit.
This object is achieved by the according to the invention, which measuring comprises a plurality of sun sensors arranged such that the measuring of the components of the sun vector is made possible in a plane in the full angular range (0.ltoreq..alpha..sub.1 .ltoreq.2.pi.) and perpendicularly to this plane in a limited angular range .vertline..alpha..sub.2 .vertline..ltoreq..alpha..sub.2max (.alpha..sub.2max &lt;90.degree.). This is based on a satellite-fixed system of coordinates, preferably a cartesian XYZ-system with the roll axis X, the pitch axis Y as well as the yaw axis Z, in which case, in the geosynchronous orbit, the roll axis is generally oriented in the orbiting direction, the yaw axis is oriented toward the center of the earth, and the pitch axis is oriented perpendicularly to the plane of the orbit. The plane, which in the following will also be called the measuring plane and which is to be fully detected by the measuring range of the sun sensors, may, for example, be the XZ-plane. Perpendicularly to it, that is, in the direction of the Y-axis, the measuring range of the sun sensors will then be limited (.+-..alpha..sub.2max). The thus defined measuring range of the sun sensors may be formed by adjoining or overlapping fields of view of several individual sun sensors with differently aligned optical axes.
As a further essential characteristic, it is required that, instead of the speed gyroscope package, which measures in three-axes in the case of known satellites, an integrating speed gyroscope must now be provided which measures in only one axis. Its measuring axis, which is represented by a unit vector G, together with the measuring plane of the sun sensors, must now enclose an angle .beta. which is larger than (.pi./2)-.alpha..sub.2max. Therefore, the measuring axis of the speed gyroscope must be within a double cone centered in the origin 0 of the system of coordinates, the axis of symmetry of the double cone being situated perpendicularly on the measuring plane of the sun sensors. When the XZ-plane is used as the measuring plane of the sun sensors, this axis of symmetry is the Y-axis.
In a further development of the invention, it is provided that the permissible solid-angle range for the measuring axis of the speed gyroscope is limited even further. The reasons for the requirement with respect to the permissible alignment of the measuring axis of the speed gyroscope will explained in detail hereinafter.
According to another feature of the invention evaluation processes are provided to determine the rotational speed of the satellite as well as the deviation of the satellite with respect to a preset reference direction, using the measuring arrangement according to the invention A regulating system is also provided to implement attitude control maneuvers while using the measuring arrangement according to the invention as well as the above-mentioned corresponding evaluation processes. Further embodiments provide regulating processes, particularly regulator equations for use in the above-mentioned regulating system.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.